This invention relates generally to gas turbine engines and more particularly to the configuration of turbine airfoils within such engines.
In a gas turbine engine, air is pressurized in a compressor and subsequently mixed with fuel and burned in a combustor to generate combustion gases. One or more turbines downstream of the combustor extract energy from the combustion gases to drive the compressor, as well as a fan, shaft, propeller, or other mechanical load. Each turbine comprises one or more rotors each comprising a disk carrying an array of turbine blades or buckets. A stationary nozzle comprising an array of stator vanes having radially outer and inner endwalls in the form of annular bands is disposed upstream of each rotor, and serves to optimally direct the flow of combustion gases into the rotor. Collectively each nozzle and the downstream rotor is referred to as a “stage” of the turbine.
The complex three-dimensional (3D) configuration of the vane and blade airfoils is tailored for maximizing efficiency of operation, and varies radially in span along the airfoils as well as axially along the chords of the airfoils between the leading and trailing edges. Accordingly, the velocity and pressure distributions of the combustion gases over the airfoil surfaces as well as within the corresponding flow passages also vary.
Undesirable pressure losses in the combustion gas flowpaths correspond with undesirable reduction in overall turbine efficiency. One common source of turbine pressure losses is the formation of horseshoe vortices generated as the combustion gases are split in their travel around the airfoil leading edges. A total pressure gradient is effected in the boundary layer flow at the junction of the leading edge and endwalls of the airfoil. This pressure gradient at the airfoil leading edges forms a pair of counterrotating horseshoe vortices which travel downstream on the opposite sides of each airfoil near the endwall. Migration of the horseshoe vortices generates a cross-passage vortex. The horseshoe and passage vortices create a total pressure loss and a corresponding reduction in turbine efficiency. These vortices also create turbulence and increase undesirable heating of the endwalls.
It is known to use 3D contouring of the endwalls (e.g. platform or shroud) of turbine airfoils to endwall contouring design reduces the strength of the horseshoe and passage vortices and the associated pressure losses, and thereby improve the turbine efficiency.
It is further known to orient or “clock” an upstream row of turbine vanes with a downstream row of turbine vanes in order to cause the wakes from the upstream vanes trailing edges to impinges on the downstream vane leading edges, where a set of rotating blades are positioned between the two rows of vanes. This concept attempts to have the lower momentum wakes impinging on the downstream vane leading edges to keep the wakes within the boundary layers of the vanes and thereby minimize the undesirable pressure losses.
Because the wakes are chopped by the rotating blade row before reaching the downstream nozzle vane leading edges, the position of the wakes are shifted as function of the blade rotating speed. For a constant rotating RPM, the tangential speed varies from the blade root to the tip. Therefore, the wake positions are shifted non-uniformly from the hub to the tip.
Accordingly, it is desirable to minimize vortex effects while also providing better alignment of nozzle wakes with a downstream nozzle.